Gas turbine component with cooling aperture having shaped inlet and method of forming the same

ABSTRACT

A method of manufacturing a cooled gas turbine component includes forming a core with an outer surface. The outer surface includes a core feature. The method also includes casting an outer wall of an airfoil about the core. The outer wall has an exterior surface and an interior surface. The interior surface includes a shaped inlet portion that corresponds to the core feature. Moreover, the method includes forming an outlet portion through the outer wall to fluidly connect the outlet portion to the shaped inlet portion. The shaped inlet portion and the outlet portion cooperatively define a cooling aperture through the outer wall.

CROSS REFERENCE TO RELATED APPLICATION

This application is a divisional of U.S. patent application Ser. No.15/266,481, filed Sep. 15, 2016, the entire disclosure of which isincorporated by reference herein.

TECHNICAL FIELD

The present disclosure generally relates to a gas turbine component and,more particularly, relates to a gas turbine component with a coolingaperture having a shaped inlet and a method of forming the same.

BACKGROUND

Gas turbine engines are generally used in a wide range of applications,such as aircraft engines and auxiliary power units. In a gas turbineengine, air is compressed in a compressor, mixed with fuel, and ignitedin a combustor to generate hot combustion gases, which flow downstreaminto a turbine section. In a typical configuration, the turbine sectionincludes airfoils, such as stator vanes and rotor blades, disposed in analternating sequence along the axial length of a generally annular hotgas flow path. The rotor blades are mounted at the periphery of one ormore rotor disks that are coupled in turn to a main engine shaft. Hotcombustion gases are delivered from the engine combustor to the annularhot gas flow path, thus resulting in rotary driving of the rotor disksto provide an engine output.

Due to the high temperatures in many gas turbine engine applications, itis desirable to regulate the operating temperature of certain enginecomponents, particularly those within the mainstream hot gas flow pathin order to prevent overheating and potential mechanical issuesattributable thereto. As such, it is desirable to cool the airfoils ofthe rotor blades and stator vanes to prevent or reduce oxidation,thermo-mechanical fatigue, and/or other adverse impacts to the airfoil.Mechanisms for cooling turbine airfoils include ducting cooling airthrough internal passages and then venting the cooling air through holesformed in the airfoil. However, given the high temperature of engineoperation, cooling remains a challenge.

Accordingly, it is desirable to provide gas turbine engines withimproved airfoil cooling. Furthermore, other desirable features andcharacteristics of the present invention will become apparent from thesubsequent detailed description of the invention and the appendedclaims, taken in conjunction with the accompanying drawings and thisbackground of the invention.

BRIEF SUMMARY

In one embodiment, a method of manufacturing a gas turbine component fora gas turbine engine is disclosed. The method includes forming a corewith an outer surface. The outer surface includes a core feature. Themethod also includes casting an outer wall of an airfoil about the core.The outer wall has an exterior surface and an interior surface. Theinterior surface includes a shaped inlet portion that corresponds to thecore feature. Moreover, the method includes forming an outlet portionthrough the outer wall to fluidly connect the outlet portion to theshaped inlet portion. The shaped inlet portion and the outlet portioncooperatively define a cooling aperture through the outer wall.

In another embodiment, a cooled gas turbine component for a gas turbineengine is disclosed. The gas turbine component includes an airfoil.Also, the gas turbine component includes an outer wall of the airfoil.The outer wall has an exterior surface and an interior surface. Also,the gas turbine component includes a cooling aperture that extendsthrough the outer wall. The cooling aperture includes a cast inletportion included on the interior surface. The cooling aperture alsoincludes an outlet portion extending through the outer wall and fluidlyconnected to the inlet portion. The inlet portion has a width and adepth, and the width of the inlet portion gradually reduces along thedepth of the inlet portion toward the outlet portion.

BRIEF DESCRIPTION OF THE DRAWINGS

The present disclosure will hereinafter be described in conjunction withthe following drawing figures, wherein like numerals denote likeelements, and wherein:

FIG. 1 is a schematic side view of a gas turbine engine according toexemplary embodiments of the present disclosure;

FIG. 2 is a perspective view of a gas turbine component of the gasturbine engine according to exemplary embodiments of the presentdisclosure;

FIG. 3 is a section view of an airfoil of the gas turbine component ofFIG. 2;

FIG. 4A is a detail view of an outer wall and a cooling aperture of theairfoil of FIG. 3 according to an example embodiment;

FIG. 4B is a detail view of the outer wall and the cooling aperture ofthe airfoil of FIG. 3 according to an additional example embodiment;

FIG. 5 is a flowchart illustrating a method of manufacturing the gasturbine component of FIG. 2;

FIGS. 6-20B are schematic views illustrating the method of FIG. 5; and

FIG. 21 is a detail view of the outer wall and cooling aperture of theairfoil according to additional embodiments of the present disclosure.

DETAILED DESCRIPTION

The following detailed description is merely exemplary in nature and isnot intended to limit the present disclosure or the application and usesof the present disclosure. Furthermore, there is no intention to bebound by any theory presented in the preceding background or thefollowing detailed description.

Broadly, exemplary embodiments disclosed herein include gas turbineengines with turbine components having improved cooling characteristics.Methods of manufacturing the turbine components are also disclosed. Inparticular, exemplary embodiments include turbine airfoils with an outerwall having at least one cooling aperture with a shaped inlet. Theshaped inlet may increase flow rate through the cooling aperture forimproved cooling. Additionally, methods of manufacturing the turbinecomponent may include casting the shaped inlet and subsequently formingan outlet of the cooling aperture through the casting to fluidly connectthe outlet to the shaped inlet. Additionally, in some embodiments, thecooling aperture may be adjusted (e.g., by widening the hole and/orchanging the axis of the outlet) to change flow characteristics of thecooling aperture. Other details of the present disclosure will bediscussed below.

FIG. 1 is a cross-sectional view of a gas turbine engine 100 accordingto an exemplary embodiment. Although FIG. 1 depicts a turbofan engine,in general, exemplary embodiments discussed herein may be applicable toanother type of engine without departing from the scope of the presentdisclosure. The gas turbine engine 100 may form part of, for example, anauxiliary power unit for an aircraft or a propulsion system for anaircraft. However, the gas turbine engine 100 may be included on anothervehicle without departing from the scope of the present disclosure.Instead of being included on a vehicle, the gas turbine engine 100 mayalso be supported by a stationary mount in some embodiments.

The gas turbine engine 100 has an overall construction and operationthat is generally understood by persons skilled in the art. The gasturbine engine 100 may be disposed in an engine case 101 and may includea fan section 120, a compressor section 130, a combustion section 140, aturbine section 150, and an exhaust section 160, which are arrangedsequentially along a longitudinal axis 180. As used herein, the term“axial” refers to a direction generally parallel to the longitudinalaxis 180. A radial axis 190 is also included in FIG. 1 for referencepurposes. The term “radial” as used herein refers to a directiongenerally parallel to the radial axis 190 and perpendicular to thelongitudinal axis 180.

The fan section 120 may include a fan, which draws in and acceleratesair. A fraction of the accelerated air from the fan section 120 isdirected through a bypass section 170 to provide a forward thrust. Theremaining fraction of air exhausted from the fan is directed into thecompressor section 130.

The compressor section 130 may include a series of compressors thatraise the pressure of the air directed into it from the fan section 120.The compressors may direct the compressed air into the combustionsection 140.

In the combustion section 140, the high pressure air is mixed with fueland combusted. The combusted air is then directed into the turbinesection 150.

The turbine section 150 may include a series of rotor assemblies 192 andstator assemblies 194, both of which are represented schematically inFIG. 1. The combusted air from the combustion section 140 expandsthrough the rotor and stator assemblies 192, 194 and causes the rotorassemblies 192 to rotate a main engine shaft for energy extraction. Theair is then exhausted through a propulsion nozzle disposed in theexhaust section 160 to provide additional forward thrust.

Within the turbine section 150, the rotor assemblies 192 may include aplurality of rotor blades 200, an example embodiment of which isillustrated in FIG. 2. As is known, the rotor blade 200 may be mountedon a rotor disc, which in turn is coupled to the engine shaft. A turbinestator directs the air toward the rotor blade 200. The air impinges uponthe rotor blade 200, thereby driving the rotor assembly 192 for powerextraction.

To allow the turbine section 150 to operate at desirable elevatedtemperatures, certain components are cooled. For example, in someembodiments, the rotor blade 200 may include cooling apertures thatinclude one or more features of the present disclosure. In additionalembodiments, one or more of the stator assemblies 194 (e.g., an airfoilof the stator assembly 194) may include the cooling apertures of thepresent disclosure. As will be discussed, the cooling aperture mayinclude a shaped inlet portion. The shaped inlet portion providesincreased flow for improved cooling. Manufacturing techniques are alsodiscussed below for providing the shaped inlet portion to one or more ofthese turbine components.

FIG. 2 illustrates an exemplary aircraft jet engine turbine rotor blade200. It will be appreciated that the rotor blade 200 may have alternateconfigurations or arrangements without departing from the scope of thepresent disclosure.

The rotor blade 200 includes an airfoil 202, a platform 204, and a root206. The platform 204 is configured to radially contain turbine airflowwithin a shroud of the turbine section 150. The root 206 extends fromthe underside of the platform 204 and is configured to couple the blade200 to a turbine rotor disc (not shown). In general, the rotor blade 200may be made from any suitable material, including high heat and highstress resistant aerospace alloys, such as nickel based alloys,Mar-M-247, single crystal materials, directionally solidified materials,or others.

The airfoil 202 projects radially outwardly from the platform 204 andterminates at a blade tip 220. The airfoil 202 is formed by a body 208with an outer wall 209. The outer wall 209 may include a first portion210 and a second portion 212 that cooperate to define an airfoil shape.The first portion 210 of the outer wall 209 defines a pressure side witha generally concave shape, and the second portion 212 of the outer wall209 defines a suction side with a generally convex shape. In a chordwisedirection, the portions 210, 212 of the outer wall 209 are joined at aleading edge 214 and trailing edge 216. As used herein, the term“chordwise” refers to a generally longitudinal dimension along theairfoil 202 from the leading edge 214 to the trailing edge 216.

As noted above, the rotor blade 200, particularly the airfoil 202, maybe subject to extremely high temperatures resulting from high velocityhot gases ducted from the combustion section 140 (FIG. 1). Ifunaddressed, the extreme heat may affect the useful life of an airfoiland/or impact the maximum operating temperature of the engine.Accordingly, cooling is provided for the rotor blade 200 (e.g., to theairfoil 202) to maintain blade temperature at an acceptable level, asdescribed in greater detail below. Such cooling may include an internalcooling system 221 that directs cooling air from the compressor section130 into inlets in the root 206 and through internal cavities andpassages of the blade 200 to cool the airfoil 202 via convection andconduction. The air flowing through the internal cooling system 221 mayflow out of the airfoil 202 through one or more cooling apertures 222,which are shown in FIG. 2.

The cooling apertures 222 may extend through the outer wall 209. It willbe appreciated that the number of cooling apertures 222 and thearrangement of the cooling apertures 222 may vary without departing fromthe scope of the present disclosure. The cooling apertures 222 may berelatively small and arranged generally in rows and/or columns on theairfoil 202, proximate the leading edge 214, the blade tip 220, and/orother areas of the airfoil 202. As shown in FIG. 2, the coolingapertures 222 may be substantially circular. The cooling apertures 222may provide cooling airflow to the leading edge 214, the blade tip 220,and/or to areas of the blade 202 between the leading and trailing edges216.

In some embodiments, the cooling apertures 222 may provide film coolingto the blade 200. Specifically, the cooling apertures 222 may bearranged to provide a cooling film of fluid onto the exterior surface(i.e., the hot side surface) of the airfoil 202.

FIG. 3 is a cross-sectional view of the airfoil 202 of FIG. 2 inaccordance with an exemplary embodiment. As shown, the cross-sectionalview may generally correspond to a cross-sectional view through anaxial-chordwise plane. It will be appreciated that the cross section ofFIG. 3 may represent a simplified embodiment and that the airfoil 202may include additional features that are not illustrated.

As discussed above, the airfoil 202 may include the outer wall 209 withfirst and second portions 210, 212 joined at the leading edge 214 andthe trailing edge 216. As shown in FIG. 3, the outer wall 209 includesan exterior surface 228 and an interior surface 230. The outer wall 209may have a thickness 232, which is measured between the exterior surface228 and the interior surface 230. In some embodiments, the outer wall209 may define a shell-like outer periphery of the airfoil 202.

Also, as noted above, the airfoil 202 may include the cooling system221, which cools the airfoil 202. Various features of the cooling system221 will be described in detail below according to exemplaryembodiments.

It will be appreciated that, although the cooling system 221 is shownand described in relation to the airfoil 202 of the rotor blade 200, oneor more features of the cooling system 221 may be incorporated withinanother area of the rotor blade 200 and/or within another turbinecomponent. For example, the cooling system 221 may be incorporatedwithin a stator assembly 194 of the turbine section 150 of the engine100 without departing from the scope of the present disclosure.

In exemplary embodiments, the cooling system 221 may form part of a highefficiency, multi-walled turbine airfoil cooling arrangement or aserpentine airfoil cooling arrangement. Other cooling systems 221 may beprovided, including those of different wall and cavity structures.

The cooling system 221 may include one or more internal voids 234defined within the airfoil 202. The internal void 234 may comprise acavity, a passageway, a channel, a pocket, a hollow, or other voidwithin the airfoil 202. As shown in the embodiment of FIG. 3, theinternal void 234 may include a forward cavity 236, a rearward cavity238, and an internal channel 240. An internal wall 242 may separate theforward cavity 236 and the rearward cavity 238. The internal channel 240may fluidly connect the forward and rearward cavities 236, 238. In someembodiments, the internal channel 240 may also be directed toward apredetermined area of the interior surface 230 for impingement coolingpurposes. The internal channel 240 may have a width (e.g., a diameter),which is indicated at 244.

The cooling system 221 may additionally include the cooling apertures222 mentioned above. One of the cooling apertures 222 is included inFIG. 3 and is shown in detail in FIG. 4A. The cooling aperture 222 mayextend through the second portion 212 of the outer wall 209, between theexterior surface 228 and the interior surface 230. An outer end 246 ofthe cooling aperture 222 may be disposed proximate the exterior surface228 of the outer wall 209. An inner end 248 of the cooling aperture 222may be disposed proximate the interior surface 230 of the outer wall209.

The cooling aperture 222 may be in fluid communication with the internalvoid 234 in some embodiments. For example, the inner end 248 may befluidly connected and open to the forward cavity 236 in someembodiments. Thus, cooling fluid may flow from the forward cavity 236,into the inner end 248 of the cooling aperture, and out of the airfoil202 via the outer end 246. Thus, the inner end 248 may define an inletof the cooling aperture 222, and the outer end 246 may define an outletof the cooling aperture 222.

As shown in FIG. 4A, the cooling aperture 222 may extend generally alongan axis 252. In some embodiments, the axis 252 may be substantiallystraight. In some embodiments, the axis 252 of the aperture 222 may becanted (i.e., disposed at an angle) relative to the exterior surface 228of the airfoil 202. In other words, as shown in FIG. 4A, the axis 252 ofthe aperture 222 may be disposed at a positive, acute angle 260 relativeto an intersecting imaginary line 258 that extends normal to theadjacent exterior surface 228 of the airfoil 202. The angle 260 may bechosen according to a desired film cooling airflow delivered to theexterior surface 228. For example, in the embodiment illustrated, theangle 260 allows cooling air to exit the cooling aperture 222 and flowgenerally along the exterior surface 228, toward the trailing edge 216.In some embodiments, the axis 252 may be between approximately zero andseventy-eight degrees)(0°-78° relative to the line 258. Also, the axis252 may be disposed at any angle relative to the gas flow direction(e.g., typically between zero and ninety degrees)(0°-90° from thedownstream direction).

The cooling aperture 222 may have a length 254 that is measured along(e.g., parallel to) the axis 252. The length 254 may be measured fromthe outer end 246 to the inner end 248 of the cooling aperture 222.

The cooling aperture 222 may also have a width 256 (e.g., diameter). Thewidth 256 may be measured transverse to (e.g., perpendicular to) theaxis 252 between opposing areas of an inner surface 250 of the aperture222.

There may be several differences between the cooling apertures 222 andother features of the cooling system 221. Specifically, the coolingapertures 222 may extend through the exterior surface 228 of the airfoil202, whereas the forward cavity 236, rear cavity 238, and internalchannel 240 are enclosed within the airfoil 202. Also, the width 256 ofthe cooling apertures 222 may be substantially smaller than the width244 of the internal channel 240. Specifically, whereas the internalchannel 240 may have a width 244 of at least 0.040 inches, the coolingaperture 222 may have a width 256 between approximately 0.012 to 0.030inches. Also, the cooling apertures 222 may provide cooling (e.g., filmcooling) to the exterior surface 228, whereas the forward cavity 236,rear cavity 38, and/or internal channel 240 may provide cooling to theinterior surface 230.

As shown in FIG. 4A, the cooling aperture 222 may be subdivided intovarious portions along its length 254. For example, the cooling aperture222 may include an inlet portion 262 and an outlet portion 264. Theinlet and outlet portions 262, 264 may be in fluid communication witheach other. Also, the outlet portion 264 may intersect the inlet portion262 at an intersection 266. The intersection 266 may extend continuouslyand annularly about the axis 252. In some embodiments, the intersection266 may be substantially circular. In other embodiments, theintersection 266 may be substantially ovate.

As shown in FIG. 4A, the outlet portion 264 may be defined by an outletsurface 268. The outlet surface 268 may extend along the axis 252between the intersection 266 and an outer rim 267 proximate the exteriorsurface 228 of the airfoil 202. In some embodiments, the width 256(i.e., diameter) of the outlet portion 264 may remain substantiallyconstant along its length from the intersection 266 to the exteriorsurface 228. Also, the outlet surface 268 may be centered about the axis252 in some embodiments. Thus, as explained above, the outlet portion264 may be disposed at an angle 260 relative to the exterior surface 228of the airfoil 202.

The inlet portion 262 may be defined by a concave inlet surface 270. Theinlet surface 270 may extend along the axis 252 between an inner rim 269and the intersection 266. In some embodiments, the width 256 of theinlet portion 262 may vary along its length from the inner rim 269 tothe intersection 266. For example, the width 256 may gradually reduce(e.g., taper) from the inner rim 269 to the intersection 266. Stateddifferently, the width 256 of the inlet portion 262 at the intersection266 may be substantially equal to the width 256 of the outlet portion264 at the intersection as shown in FIG. 4A. In other embodiments thatwill be discussed below, the width 256 of the inlet portion 262 at theintersection 266 may be greater than the width 256 of the outlet portion264 at the intersection 266. In this latter embodiment, a step may bedefined at the transition 266. Also, in some embodiments, at least partof the concave inlet surface 270 may face generally toward the central,interior region of the airfoil 202. As such, the concave inlet surface270 of the inlet portion 262 may be considered part of the interiorsurface 230 of the outer wall 209.

The inlet surface 270 may be defined according to a variety of shapes.For example, the inlet surface 270 may be at least partially conic(e.g., frustoconic) in some embodiments. As such, the inlet surface 270may have two-dimensional curvature. In other embodiments, the inletportion 262 may be concave and generally dome-shaped, hemispherical, orotherwise rounded. As such, the inlet surface 270 may havethree-dimensional curvature.

In some embodiments, the frustoconic inlet portion 262 may besubstantially centered about the axis 252. Thus, as shown in FIG. 4A,the axis 252 of the inlet portion 262 may be disposed at an angle 260relative to the adjacent areas of the interior surface 230 of the outerwall 209. Accordingly, the inlet portion 262 and the outlet portion 264may be canted generally toward the trailing edge 216 of the airfoil 202.

FIG. 4B shows additional embodiments of the inlet portion 262′ of thecooling aperture 222′. As shown, the inlet surface 270′ may be contouredconvexly in cross section. The inlet surface 270′ may have a radius 271′that provides a smooth transition from adjacent areas of the interiorsurface 230′ to the outlet surface 268′. Also, as shown in FIG. 4B, theinlet portion 262′ may be centered about an axis 253′. The axis 253′ maybe disposed at an angle 273′ relative to the axis 252′ of the outletportion 264′. The axis 273′ may intersect the axis 252′ proximate theintersection 266′ in some embodiments. Also, while the axis 273′ isshown as a straight axis in FIG. 4B, the axis 273′ may be curved inother embodiments. It will be appreciated that the axis 273′ and othercharacteristics of the inlet portion 262′ may be chosen according todesired flow characteristics for the cooling aperture 222′. For theremaining discussion, the embodiment of FIG. 4A will be discussed.However, it will be appreciated that the disclosure may relate to theembodiment of FIG. 4B as well.

In some embodiments, the outlet portion 264 may comprise the majority ofthe length 254 of the cooling aperture 222. Also, the length of theoutlet portion 264 (measured from the outer rim 267 to the intersection266 in FIG. 4A) may be at least 1× the width 256 of outlet portion 264.For example, the depth 274 of the inlet portion 262 (measured, forexample, between the inner rim 269 and the intersection 266) may be, atmost, one third (⅓) of the length 254 of the cooling aperture 222. Insome embodiments, a ratio of the depth 274 of the inlet portion 262 tothe length of the outlet portion 264 may be approximately 0.1-0.5 foroutlet portions 264 that are under 0.5 inches in diameter. The ratio ofthe depth 274 to the length of the outlet portion 264 may be 0.05-0.4for outlet portions 264 that are larger than 0.5 inches in diameter.Additionally, in embodiments like those illustrated in FIG. 4A, thewidth 256 of the inlet and outlet portions 262, 264 may be substantiallyequal at the intersection 266. In other embodiments, the width 256 ofthe inlet portion 262 at the intersection 266 may be somewhat greaterthan the width 256 of the outlet portion 264 at the intersection 266. Ineither case, the width 256 of the inlet portion 262 at the intersection266 is, at most, 1.4 times (1.4×) the width 256 of the outlet portion264 at the intersection 266. These dimensions yield significant flowincreases for the cooling aperture 222, and yet the airfoil 202 mayremain quite strong and robust. Also, these dimensions ensure that theairfoil 202 can be highly manufactureable. Additionally, thesedimensions may account for tolerancing in the manufacturing process(e.g., when forming the outlet portion 264 to intersect the inletportion 262). It will be appreciated that these dimensions can apply tothe embodiment of FIG. 4B. For example, the ratio of the radius 271′ tothe length of the outlet portion 264′ may be approximately 0.1-0.5 foroutlet portions 264′ that are under 0.5 inches in diameter. Also, theratio of the radius 271′ to the length of the outlet portion 264′ may be0.05-0.4 for outlet portions 264′ that are larger than 0.5 inches indiameter. Also, in the embodiment of FIG. 4B, the ratio of the radius271′ to the width 256′ of the outlet portion 264′ may be approximately0.05 to 0.5.

The inlet portion 262 may be open to the forward cavity 236 within theairfoil 202. The outlet portion 264 may be in fluid communication withthe inlet portion 262. Therefore, a fluid flowpath may be defined fromthe forward cavity 236, to the inlet portion 262, and out of the airfoil202 via the outlet portion 264 for film cooling of the airfoil 202.

The inlet portion 262 may define a widened and shaped inlet portion ofthe cooling aperture 222. The inlet portion 262 may provide a gradualtransition along the interior surface 230 to the outlet portion 264. Insome embodiments, the inlet portion 262 may define a chamfer orchamfer-like feature for the outlet portion 264 of the cooling aperture222. Thus, the profile of the inlet surface 270, the tapering width 256of the inlet portion 262, and/or other features may increase fluid flowthrough the cooling aperture 222. As a result, the airfoil 202 may becooled more efficiently and effectively. Also, because the inlet portion262 increases fluid flow, the width 256 of the outlet portion 264 may bereduced, making the outer wall 209 stronger and more robust.

Methods of manufacturing the airfoil 202 will now be discussed withreference to FIGS. 5-20B. An example embodiment of the method 300 isillustrated in FIG. 5. The method 300 is illustrated schematically inFIGS. 6-20B as will be discussed.

As shown in FIG. 5, the method 300 may begin at 302, in which a die(i.e., a “core die”) is formed, such as the die 400 illustratedschematically in FIG. 6. As shown, the die 400 may include a firstmember 402 and a second member 404 that cooperate to define an internalcavity 406. In other words, the first member 402 and second member 404may define an internal surface 408, and the internal surface 408 maydefine an internal cavity 406 of the die 400.

As shown in FIGS. 6 and 7, the internal surface 408 is shaped to includea core die feature 410. The core die feature 410 may be a concavefeature that corresponds substantially to the inlet portion 262 of thecooling aperture 222 of FIGS. 3 and 4. The core die feature 410 may beshaped substantially the same as the inlet portion 262 of the coolingaperture 222 described above. Thus, the core die feature 410 may be arecess, void, pocket, or other concave feature in the internal surface408. It will be understood that the inlet portion 262 of the coolingaperture 222 may ultimately correspond in shape to the core die feature410.

Next, the method 300 of FIG. 5 may continue at 304, in which a core 414is formed within the internal cavity 406 of the core die 400. As shownin FIG. 8, a core material 412, such as a ceramic material, may beprovided to the internal cavity 406. The core material 412 may be moldedwithin the internal cavity 406 and cured to form the core 414. An outersurface 415 of the core 414 may correspond to the internal surface 408of the internal cavity 406 of the core die 400. Thus, the core outersurface 415 may include a core feature 416. The core feature 416 may bea projection that corresponds in shape substantially to the core diefeature 410. More specifically, the core feature 416 may be inverselyshaped according to the core die feature 410. Then, as shown in FIG. 9,the core 414 may be removed from the core die 400.

It will be appreciated that the core 414 may be formed in other ways aswell. For example, additive manufacturing techniques (e.g., 3D printing)may be employed for forming the core 414. These techniques may also beused to form the core feature 416 of the core 414.

Subsequently, the method 300 may continue at 306, in which a tool 418 isfabricated. As shown in FIG. 10, the tool 418 may include a first member420 and a second member 422, which cooperate to define an internalcavity 424. In other words, the first and second members 422 may includerespective inner surfaces 426 that define the cavity 424. The innersurfaces 426 may correspond, in some embodiments, to the exteriorsurface 228 of the airfoil 202. Thus, as will be discussed, the exteriorsurface 228 may be formed substantially according to the inner surfaces426.

Next, at 308, the core 414 may be disposed within the cavity 424 asshown in FIG. 10. The core 414 may be supported within the cavity 424 soas to define a region 428 between the outer surface 415 of the core 414and the internal surface 426 of the tool 418.

Subsequently, at 310 of the method 300, a shell material 429, such aswax, may be provided within the region 428. The shell material 429 maybe hardened to define a shell 430 about the core 414. Accordingly, afirst intermediate article 432 may be formed that includes the core 414and the shell 430.

The method 300 may continue at 311. At 311, the first intermediatearticle 432 may be removed from the cavity 424 of the tool 418, and thefirst intermediate article 432 may be dipped one or more times in aslurry material 434. The slurry material 434 may be a ceramic material.As shown in FIG. 13, the slurry material 434 may harden about and encasethe first intermediate article 432 in a second shell 436 to define asecond intermediate article 438, which includes the second shell 436,the first shell 430, and the core 414. Subsequently, as shown in FIG.14, the first shell material 429 may be removed. For example, the firstshell material 429 may be melted out. Once the first shell material 429is removed, a third intermediate article 440 may be formed with an openregion 442 defined between the outer surface 415 of the core 414 and aninternal surface 443 of the shell 436.

Then, at 312 of the method 300, the airfoil 202 may be cast about thecore 414 and within the region 442. Specifically, as shown in FIG. 15,an airfoil material 444 may be provided to the region 442. The airfoilmaterial 444 may be cured to form the outer wall 209 and other portionsof the airfoil 202. As shown in FIG. 15, the interior surface 230 of theouter wall 209 may be cast and formed according to the outer surface 415of the core 414 such that the interior surface 230 includes the inverseof the core feature 416. Accordingly, the inlet portion 262 is formed onthe interior surface 230 of the outer wall 209 according to the corefeature 416.

Then, as shown in FIG. 16, the second shell 436 may be removed, forexample, by breaking the second shell 436 to reveal the outer wall 209.Next, at 314 of the method 300, the core 414 may be removed. Forexample, the core material 412 may be chemically leached from within theouter wall 209, leaving the hollow airfoil 202.

Subsequently, at 316 of the method 300, the outlet portion 264 of thecooling aperture 222 may be formed. For example, as shown in FIGS. 18and 19, the outlet portion 264 may be formed by EDM, drilling, laserdrilling, or other machining methods. A cutting tool 446, such as awire, drill bit, laser cutting tool, etc., may progressively removematerial through the outer wall 209 in a direction from the exteriorsurface 228 toward the interior surface 230 to form the outlet portion264. The outlet portion 264 may be cut so that it intersects with theinlet portion 262 of the cooling aperture 222 at the intersection 266.Once the outlet portion 264 is formed, what remains may be the airfoil202 with the cooling aperture 222 shown in FIGS. 3, 4A, and/or 4B. Thecooling aperture 222 will include the outlet portion 264 as well as thegradually widening inlet portion 262.

During the development process of a cooled vane or blade, theorientation and diameter of the cooling aperture 222 may change based ontest results. There may also be tolerances on where the intersection 266is located on the inner surface 230 (i.e., where the outlet portion 264breaks out on the inner surface 230). Accordingly, in some embodiments(for example, in which the airfoil 202 is to be tested before in-flightuse), the method 300 may continue at 318. At 318, the airfoil 202 may betested, for example, to determine the flow and cooling characteristicsof the cooling aperture 222. In some embodiments, the airfoil 202 may betested in connection with CFD modeling techniques and tools.

Next, the method 300 may continue to 320, wherein it determined whetherto adjust the cooling aperture 222 to provide more desirable flow andcooling characteristics. If no adjustments are needed, the method 300may finish. If, on the other hand, the testing of 318 indicates thatadjustments are needed for the cooling aperture 222, then the method maycontinue to 322.

At 322, it may be determined whether it is necessary to re-cast theairfoil 202. In some cases, such as the embodiment of FIG. 20A, thecooling aperture 222 may be adjusted without the need to re-cast theairfoil 202 (i.e., 322 answered negatively). Thus, at 324 of the method300, the outlet portion 264 of the cooling aperture 222 formed at 316(and illustrated in FIG. 19) is re-formed, re-machined, reamed,re-drilled, re-conditioned, or otherwise adjusted. Specifically, in FIG.20A, the original width 256 of the outlet portion 264 is increased to alarger width 256′. In some embodiments, a larger diameter cutting tool446′ (larger than the cutting tool 446 of FIG. 19) may be used to widenthe outlet portion 264. Accordingly, the transition between the inletand outlet portions 262, 264 may be made to be more gradual. Forexample, a sharp corner and/or step at the intersection of the inlet andoutlet portions 262, 264 may be removed. This may improve flowcharacteristics through the cooling aperture 222. After adjustments aremade at 324, the method 300 may loop back to the testing of 318 and theadjustment inquiry of 320. If more adjustments are needed, then themethod 300 may continue to 322 and so on. If there are no furtheradjustments needed, then the method 300 may end.

In other embodiments of 322 of the method 300, it may be necessary tore-cast the airfoil 202 (i.e., 322 answered positively). Thus, themethod 300 may continue to 326, wherein the core 414 is re-formed. Thecore 414 may be re-formed using the same core die 400 formed at 302 andillustrated in FIGS. 6-8. Next, the method 300 may loop back to 308 andproceed as discussed above until the airfoil 202 is re-cast at 312. Thecore 414 may then be removed at 314. Next, at 316, the outlet portion264 may be formed in the re-cast airfoil 202″ as illustrated in FIG.20B. As shown, a cutting tool 446″ may be used to form the outletportion 264″ through the re-cast outer wall 209″. In some embodiments,the cutting tool 446″ may have a greater width (i.e., greater diameter)than the cutting tool 446 used previously (FIG. 19). Also, as shown inFIG. 20B, the outlet portion 264″ may be formed along a different axis252″. Specifically, the outlet portion 264″ may be formed to extendthrough the outer wall 209 along an axis 252″ that is disposed at anangle 452″ relative to the original axis 252. Then, the method 300 maycontinue to the testing of 318 and the adjustment inquiry of 320. Ifmore adjustments are needed, then the method 300 may continue to 322 andso on. If no further adjustments are necessary, then the method 300 end.

It will be appreciated that the adjustments shown in FIG. 20A and 20Bmay be made in a convenient and efficient way. For example, in the caseof FIG. 20A, the outlet portion 264 can be widened without having tomake a new core, new molding dies or tools, etc. Additionally, in thecase of FIG. 20B, the outlet portion 264″ may be widened and re-orientedwithout having to make a new core die 400, etc.

Referring now to FIG. 21, alternative embodiments of the presentdisclosure will now be discussed. FIG. 21 shows a cross section of theouter wall 209″. The cross section of the outer wall 209″ may besubstantially similar to the outer wall 209 shown in FIG. 4, except thatthe outlet portion 264″ has not been drilled and is only shown inphantom.

As shown in FIG. 21, the areas of the outer wall 209″ that are adjacentto the inlet portion 262″ may include a thickened area 470″. Thethickened area 470″ may provide added thickness 232 that is localizedabout the concavity of the inlet portion 262″. Accordingly, thethickened area 470″ provides added strength and reinforcement proximatethe inlet portion 262″. The thickened area 470″ may be annular and mayproject toward the central region of the wall 209″. The thickened area470″ may continuously surround the inlet portion 262″ about the axis252″.

It will be appreciated that the manufacturing method 300 discussed abovemay be used to manufacture the thickened area 470″ and other areas ofthe outer wall 209″. For example, at 312 of the method 300, the outerwall 209″ may be cast about a core 414″, and the inner surface 230″(including the thickened area 470″) may be formed according to the outersurface 415″ of the core 414″. The core 414″ may include a recess 472″,and the core feature 416″ may project outwardly from the recess 472″.Thus, the thickened area 470″ may be formed inversely according to thesurfaces of the recess 472″. Likewise, the inlet portion 262″ may beformed inversely according to the surfaces of the core feature 416″.

Accordingly, the shaped inlet portions 262 of the cooling apertures 222of the airfoil 202 provide several advantages. The shaped inlet portion262 improves the flow coefficient of the cooling aperture 222, allowingmore air to pass through a given hole width 256. For an aperture 222that is perpendicular to the wall 209, the shaped inlet portion 262 canincrease flow at least 15%, to 40%. Moreover, in cases in which theshaped inlet portion 262 extends continuously about the axis 252 of theoutlet portion 264, the flow can be less sensitive to variations in thedirection of flow and wall thickness variations. Additionally, thecooling apertures 222 can be less likely to plug with dirt or otherdebris because the inlet portion 262 is tapered instead of having asharp inlet edge. This is because flow through the cooling apertures 222is less likely to result in recirculation zones, which can causeparticle build-up. Moreover, the shaped inlet portions 262 can directand channel the air, minimizing flow separations within the aperture.Accordingly, film cooling can occur in an effective manner.

Furthermore, manufacturing of the airfoil 202 can be completed in anefficient manner. The casting and subsequent drilling or EDM methodsdescribed above can be completed in a controlled fashion for highmanufacturing accuracy. The methods provide time savings as well. Also,testing and adjusting the cooling apertures 222 may be completed in aconvenient manner because, instead of having to form another core withnew tooling, molds, etc., the outlet portion 264 can be adjusted byre-drilling as discussed above.

While at least one exemplary embodiment has been presented in theforegoing detailed description, it should be appreciated that a vastnumber of variations exist. It should also be appreciated that theexemplary embodiment or exemplary embodiments are only examples, and arenot intended to limit the scope, applicability, or configuration of thepresent disclosure in any way. Rather, the foregoing detaileddescription will provide those skilled in the art with a convenient roadmap for implementing an exemplary embodiment of the present disclosure.It is understood that various changes may be made in the function andarrangement of elements described in an exemplary embodiment withoutdeparting from the scope of the present disclosure as set forth in theappended claims.

What is claimed is:
 1. A cooled gas turbine component for a gas turbineengine comprising: an airfoil; an outer wall of the airfoil, the outerwall having an exterior surface and an interior surface; and a coolingaperture that extends through the outer wall, the cooling apertureincluding: a cast inlet portion included on the interior surface; and anoutlet portion extending through the outer wall and fluidly connected tothe inlet portion; wherein the inlet portion has a width and a depth,wherein the width of the inlet portion gradually reduces along the depthof the inlet portion toward the outlet portion.
 2. The cooled gasturbine component of claim 1, wherein the outlet portion is a holehaving a substantially constant diameter through the outer wall.
 3. Thecooled gas turbine component of claim 1, wherein the inlet portion is atleast partially conic.
 4. The cooled gas turbine component of claim 1,wherein the outlet portion extends along an axis; and wherein the axisextends at an acute angle relative to the exterior surface of the outerwall.
 5. The cooled gas turbine component of claim 1, wherein the inletportion continuously encompasses the outlet portion.
 6. The cooled gasturbine component of claim 1, further comprising a thickened area thatis adjacent the inlet portion.
 7. The cooled gas turbine component ofclaim 1, wherein the depth of the inlet portion is, at most, one thirdof a length of the cooling aperture.
 8. The cooled gas turbine componentof claim 1, wherein the inlet portion comprises a chamfer of the outletportion.
 9. The cooled gas turbine component of claim 1, wherein asurface of the inlet portion has a convex curvature.
 10. The cooled gasturbine component of claim 1, wherein the outlet portion extends along astraight axis.
 11. The cooled gas turbine component of claim 4, whereinthe acute angle is, at most, seventy-eight degrees (78°).
 12. The cooledgas turbine component of claim 1, further comprising an internal wallthat separates a first cavity from a second cavity within the airfoil;further comprising an internal channel that extends through the internalwall and that fluidly connects the first cavity and the second cavity;and wherein a first width of the cooling aperture is smaller than asecond width of the internal channel.
 13. A cooled gas turbine componentfor a gas turbine engine comprising: an airfoil; an outer wall of theairfoil, the outer wall having an exterior surface and an interiorsurface; and a cooling aperture that extends through the outer wall, thecooling aperture including: a shaped inlet portion included on theinterior surface; and an outlet portion extending through the outer walland fluidly connected to the inlet portion; wherein the inlet portionhas a width and a depth, wherein the width of the inlet portion tapersand reduces in width along the depth of the inlet portion toward theoutlet portion.
 14. The cooled gas turbine component of claim 13,wherein the cooling aperture extends through a first region of the outerwall; wherein the outer wall includes a second region that is spacedapart at a distance from the cooling aperture; wherein the exteriorsurface is smooth from the first region to the second region; andwherein a first wall thickness of the outer wall in the first region isgreater than a second wall thickness of the outer wall in the secondregion.
 15. The cooled gas turbine component of claim 13, wherein theoutlet portion is a hole having a substantially constant diameterthrough the outer wall.
 16. The cooled gas turbine component of claim13, wherein the inlet portion is frustoconic.
 17. The cooled gas turbinecomponent of claim 13, wherein the outlet portion extends along astraight axis; and wherein the axis extends at an acute angle relativeto the exterior surface of the outer wall.
 18. The cooled gas turbinecomponent of claim 13, wherein the inlet portion continuouslyencompasses the outlet portion.
 19. The cooled gas turbine component ofclaim 13, wherein the depth of the inlet portion is, at most, one thirdof a length of the cooling aperture.
 20. The cooled gas turbinecomponent of claim 1, wherein a surface of the inlet portion has aconvex curvature.